This application relates to a gas turbine engine having a gear driven fan and utilizing integrally bladed rotors in a compressor section.
Gas turbine engines are known and, typically, include a fan delivering air into a bypass duct as propulsion air and further delivering a portion of air into a core engine. The air passing into the core engine moves a compressor section where it is compressed. The compressed air is delivered into a combustor where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors driving them to rotate. The turbine rotors, in turn, rotate the compressor rotors and the fan rotor.
Historically, in one common type of gas turbine engine, a single turbine rotor drove both a lower pressure compressor and a fan rotor at a common speed. This put limits on the operation of the gas turbine engine as it would be desirable to have the turbine and the lower pressure compressor rotor rotate at a higher speeds, but the fan rotor suggested speed was limited.
Another common type of gas turbine engine utilized a separate fan drive turbine rotor, which directly drove the fan rotor. The same restrictions with regard to the speed of this fan drive turbine existed due to limitations on the speed of the fan rotor.
More recently, it has been proposed to place a gear reduction between a fan drive turbine and the fan.
The compressor rotors typically utilized in gas turbine engines, such as for use on commercial aircraft, have included compressor rotors having hubs that receive removable blades.
It is known to utilize integrally bladed rotors, wherein a hub and a plurality of compressor blades are all formed as one unit. However, such rotors have only been utilized in military applications where performance takes such priority that additional cost is of no concern.